Method of making large airplane structures

ABSTRACT

A method of making a large airplane structure from a plurality of subassemblies. The method comprises drilling coordination holes in selected components which are made for accurate assembly of the subassemblies, and that will be located on the subassemblies in a position to be used to accurately position the subassemblies relative to each other so the spatial relationships between key features of the subassemblies as defined by the coordination holes make the subassemblies self-locating and intrinsically determinate of the final contour and configuration of the large airplane structure, independent of tooling. The drilling of coordination holes in the selected components is done using an end effector carried by a precision computer controlled robot that is directed to the drilling locations using a digital dataset taken directly from original digital part definition records.

This is a division of U.S. application Ser. No. 07/964,533 filed on Oct.13, 1992, and now U.S. Pat. No. 5,560,102. This invention relates to amethod and apparatus for manufacturing large mechanical structures, andmore particularly to a method and apparatus for manufacturing panels andmajor airplane fuselage sections.

BACKGROUND OF THE INVENTION

Traditional manufacturing techniques for assembling components toproduce large mechanical structures to a specified contour traditionallyhave relied on fixtured tooling techniques utilizing floor assembly jigsand templates to locate and temporarily fasten detailed structural partstogether to locate the parts correctly relative to one another. Thistraditional tooling concept usually requires at least one primaryassembly tool for each subassembly produced, and movement of the partsfrom tool to tool for manufacturing operations as they are built up.

The tooling is intended to accurately reflect the original engineeringdesign of the product, but there are many steps between the originaldesign of the product and the final manufacture of the tool. It is notunusual that the tool as finally manufactured produces parts that areoutside of the dimensional tolerances of the original part design, and,more seriously, the tool can become out of tolerance from typical harduse it receives in the factory. Moreover, dimensional variations causedby temperature changes in the factory can produce a variation in thefinal part dimensions as produced on the tool. Also, hand drilling ofthe part on the tool produce holes that are not perfectly round when thedrill is presented to the part at a slightly nonperpendicular angle tothe part, and also when the drill is plunged into the part with a motionthat is not perfectly linear. Parts can shift out of their intendedposition when they are riveted in non-round holes, and the nonuniformhole-to-rivet interference in a non-round hole lacks the strength andfatigue durability of round holes. The tolerance buildup on the part asit is moved from tool to tool can result in significant deviation fromthe original design dimensions, particularly when the part is located onthe tool at one end of the part, forseeing all of the part variation inone direction instead of centering it over the true intended position.Finally, this traditional hard tooling is expensive, difficult to changewhen design changes are implemented and takes up a large amount offactory floor space.

These disadvantages of the use of hard tooling are inherent in theconcept and, although they can be minimized by rigorous quality controltechniques, they will always be present to some extent in themanufacture of large mechanical structures using hard tooling.

SUMMARY OF THE INVENTION

Accordingly, it is an object of this invention to provide a method ofmanufacturing large mechanical structures which is independent oftraditional hard tooling to determine the placement of the partsrelative to one another and the part contour.

Another object of the invention is to provide a method of manufacturinglarge mechanical structures using intrinsic features of the part toallow them to self locate and determine part dimensions and partcontours rather than using the traditional hard tooling concepts.

It is yet another object of this invention to provide a system formanufacturing large mechanical structures that is inherently moreaccurate than the prior art and produces structures in which the partsare consistently located on the structure within the tolerance specifiedby the engineering design.

It is yet another object of the invention to provide a system formanufacturing large mechanical structures that is faster and lessexpensive than the prior art traditional techniques and requires lessfactory space and is less dependent upon the skill of workers to produceparts within the engineering tolerances specified.

These and other objects of the invention are attained in a system usinga method that utilizes spatial relationships between key features ofdetail parts or subassemblies as represented by coordination holesdrilled into the parts and subassemblies by accurate numericallycontrolled machine tools using original numerical part definitionrecords and making the parts and subassemblies intrinsically determinateof the dimensions and contour of the assembly.

DESCRIPTION OF THE DRAWINGS

The invention and its many attendant objects and advantages will becomebetter understood upon reading the following detailed description of thepreferred embodiment in conjunction with the following drawings,wherein:

FIG. 1 is a partially exploded perspective view of a portion of anairplane fuselage constructed in accordance with this invention;

FIG. 2 is an enlarged view of a junction between a stringer, a sheartie, a stringer clip and a frame member in a fuselage section made inaccordance with this invention;

FIG. 3 is a schematic representation of a process for assembling panelsin accordance with this invention;

FIG. 4 is a fuselage panel assembly cell designated as station B in FIG.3, with the left-hand bank omitted for clarity of illustration;

FIG. 5 is a perspective view of one end of a reconfigurable fixtureshown in the cell of FIG. 4, and also showing a monument for checkingthe accuracy of the machine tool in the cell;

FIG. 6 is a perspective view of the other end of the reconfigurablefixture shown in FIG. 5 and showing the fuselage skin handling system;

FIG. 7 is an elevation of an index device for positioning a fuselageskin on the reconfigurable fixture shown in FIGS. 5 and 6;

FIG. 8 is an end view of the index device shown in FIG. 7;

FIG. 9 is a perspective view of an assembled panel showing thecoordination holes in the edges of the panel and in the stringer clips;

FIG. 10 is an enlarged detail of FIG. 9;

FIG. 11 is a perspective view of two panels being joined along one edgeusing the coordination holes to form a superpanel;

FIG. 12 is a perspective view of a fixture holding a frame member fordrilling of coordination holes by an NC machine tool;

FIG. 13 is a perspective view of an assembled quarter panel, showing theframe members installed in accordance with this invention;

FIG. 14 is an enlarged perspective view of a detail in FIG. 13, showinghow the frame and stringer clip coordination holes are aligned;

FIG. 15 is a perspective view of a floor grid held in a fixture whilecoordination holes are drilled in the ends of the cross members by an NCmachine tool;

FIG. 16 is an exploded perspective view of a fuselage lower lobe showingthe floor grid and frame connection, and also the frame and superpanelconnection;

FIG. 17 is a perspective view of an assemble lower lobe in accordancewith this invention;

FIG. 18 is a perspective view of a completely assemble fuselage sectionaccording to this invention;

FIG. 19 is an elevation of the fuselage panel assembly cell shown inFIG. 4, with a touch probe carried by the robot arm instead of amachining end effector; and

FIG. 20 is a schematic representation of the computer architecture thatcontrols the fuselage panel assembly cell shown in FIG. 4.

DESCRIPTION OF THE PREFERRED EMBODIMENT

Referring now to the drawings, wherein like reference charactersdesignate identical or corresponding parts, and more particularly toFIG. 1 thereof, a portion of an airplane fuselage section is shownhaving a skin 20 unto which is fastened by riveting a plurality ofparallel stringers 22 and a plurality of shear ties 24 along stationplanes perpendicular to the longitudinal axis of the fuselage. Aplurality of stringer clips 26 are positioned in the channel of eachstringer 22 and each stringer clip 26 has a flat surface 28 which isdesigned to lie on one of the same station planes on which the shear tiesurfaces lie. A frame member 30 having a curved contour the same as thedesired contour of the airplane fuselage is riveted to the shear tiesand the stringer clips, as shown in FIG. 2.

To ensure that the faying surfaces of the shear ties 24 and the stringerclips 26 lie within the designated tolerance limits of 0.010" from theirstation planes, and that the alignment of the station planes betweenbody panels, that is, the panel-to-panel indexing of station lines, iswithin tolerance limits, so that the frame members 30 may be fastened tothe body panels and joined in alignment without the use of shims andwithout stressing the panel, the stringers 22, the shear ties 24 and thestringer clips 26 must be fastened to the fuselage skin 20 with extremeaccuracy and consistency. The consistency enables the use of statisticalprocess control to detect a trend toward an out of tolerance conditionbefore bad parts are produced so that corrective action may be taken.Accuracy of parts manufacture insures that the airplane will cometogether perfectly with no prestressed parts and no cosmeticimperfections.

The object of this invention is to provide a method and apparatus usablein the airframe industry, as well as others, which enables themanufacturer of parts with such precision and consistency.

Turning now to FIG. 3, a schematic representation of the process formanufacturing panels is shown having five stations. In the firststation, station A, fuselage skins 20, stringers 22, shear ties 24 andstringer clips 26 are brought to a fuselage assembly cell 32 at stationB where the fuselage skins 20 are loaded on to a reconfigurable holdingfixture 34 as disclosed in U.S. Pat. No. 5,249,785, the disclosure ofwhich is incorporated herein by reference. The reconfigurable fixture 34has a plurality of headers 36 which move laterally in the fixture onslides 38 to enable skins 20 of different contour to be mounted on thefixture 34 for assembly of the stringers 22, shear ties 24, and stringerclips 26. Two index devices 40 and 40' mount to the side of two of theheaders 36 as shown in FIGS. 5 and 6.

Skins 20, which in the preferred embodiment are contoured, chemicalmilled aluminum aircraft skins, are brought from a skin storage area 42via a trolley system 44 and are lifted on to the reconfigurable fixture34 by an elevator 46 to which the trolley 44 is transferred. The skins20 are lifted by the elevator 46 and positioned so that a pair of endtabs 48 and 48' are aligned with pins 50 on the index devices 40 and40'.

The index device shown in FIGS. 6 and 7, includes three precision groundalignment pins 52 which fit into precision ground bushings set into theheaders 36. The index devices 40 are held in place by bolts 54 which canbe hand tightened by star wheels 56. After the holes in the skin indextabs 48 have been placed over the pins 50 on the index devices 40 and40', a slide 58 is slid over the skin index tabs 48 to hold the skin onthe pin 50. Vacuum is applied to suction cups 59 on the headers 36 tohold the skin firmly against the headers 36, ready for machiningoperations.

Looking back at FIGS. 3 and 4, a machine tool 60 is shown mountedbetween the two reconfigurable fixtures 34. In the preferred embodiment,the machine tool 60 is a CNC machine tool such as a Jo'Mach 16/3Bproduced by JOBS, that combines articulated movements with rigidity andaccuracy of a conventional CNC machine tool. The machine tool 60controls a robot arm 62 which operates any of a number of end effectorsfor performing machining operations on the skin 20. The end effectorsare stored in a rack 64. Each end effector is supported on a clamp inits own cubicle in the rack 64 for security.

To pick up an end effector, the robot 60 rotates about a vertical axisand travels longitudinally along its support tracks concealed under aflexible track cover 66. The machine tool 60 inserts the end of therobot arm 62 into the receptacle in the end effector and a computercontroller 68 directs the clamps in the cubicle to release the endeffector to be picked up by the robot arm of the machine tool 60. Themachine tool 60 carries the end effector back to the location of theskin 20 on the fixture 34 and commences machining operations on theskin.

The machining process after the skin 20 has been accurately located onthe fixture 34, includes drilling coordination holes in the skin 20 andin the stringers 22 mounted below the skin on the fixture 34. A numberof coordination holes are drilled, sufficient to accurately locate thestringer 22 on the skin 20 when the stringers are assembled to the skinat a separate station. A drilling and centering end effector (not shown)centers the channel of the stringer 22 on the end effector so that thehole drilled by the end effector is precisely in the center of thestringer channel.

After all of the stringers 22 have been drilled with coordination holesand corresponding coordination holes have been drilled in the skin 20,the shear ties 24 are drilled and coordination holes for the shear tiesare drilled in the fuselage skin. For this purpose, an end effector isprovided for picking up the shear ties 24, placing them against thefuselage skin 20 in the correct location, and drilling two coordinationholes simultaneously through the flange of the shear tie 24 and the skin20. The shear tie 24 is presented to the shear tie placement anddrilling end effector by a parts presenter which centers the shear tiein a presentation location so that the shear tie is always presented tothe robot at exactly the same position and orientation. In this way,when the end effector picks up the shear tie 24, it always picks it upat exactly the same position so that when it is presented to the skin20, the ends of the shear tie are precisely positioned in accordancewith the original design definition for the part.

The shear ties 24 are positioned one by one at their correct location onthe skin 20 and then returned to a numbered storage tray havingindividual slots for each shear tie. In this way, it is possible toensure that, when the panel is assembled, the correct parts are placedat the positions at which the coordination holes for those parts weredrilled.

After the shear ties have all been positioned and drilled and returnedto their storage tray, a stringer clip placement and drilling endeffector is picked up by the robot arm 62 of the machine tool 60. Thestringer clip placement and drilling end effector is disclosed in U.S.Pat. No. 5,127,139, the disclosure of which is incorporated herein byreference. The stringer clip placement and drilling end effector picksup the stringer clips 26 from the parts presenter, and places them atthe correct location in the channel of the stringers 22. A clamp in theend effector squeezes the sides of the stringer against the stringerclip and a pair of opposed drills drills through the side of thestringer and through the side walls of the stringer clip 26. Acoordination hole 69 is also drilled in the end of the stringer clip 26at the same time for a purpose which will appear presently. The endeffector then releases the stringer clip where it is held in thestringer 22 by the resilience of the stringer walls.

The final machining operation is an edge routing which is performed by ahigh speed routing end effector. The machine tool 60 returns thestringer clip placement and drilling end effector to its cubicle in therack 64 and picks up the routing end effector. The machine tool 60returns to its position opposite the fixture 34 and routes the edges ofthe fuselage skin 20 to the correct dimension specified by the originalpart definition data base by accurately locating the edges relative tothe coordination holes in the skin. At the conclusion of the routing,the fuselage skin 20 conforms closely to the original engineering partdefinition. In practice, the tolerance has been in the order of lessthan 0.005" which, when the parts are assembled on the skin, produces apanel which can be assembled with extremely close conformance to theoriginal product specification without the use of shims and withoutprestressing any of the parts.

The skin 20 and the stringers 22 are removed from the fixture 34 on thetrolley 44 are deburred to remove burrs around the drilled coordinationholes, and are carried to Station C in FIG. 3 which are simple hangingracks on which the skin 20 can be hung while the shear ties 24, whichhave also been deburred to remove burrs around the drilled coordinationholes, and stringers 22 are sealed and tack fastened to the skin 20through their aligned coordination holes. The assembled panel is nowtaken to an automatic riveting machine such as a Drivematic made byGEMCOR, the shear ties 24 and stringers 22 are clamped to the skin 20,and the shear ties and stringers are drilled and riveted to the skin.

The panel 70 is next taken to Station E in FIG. 3 where the stringerclips 26 are inserted at the correct location and are held in placewhile additional holes are drilled and rivets are inserted and upset.This process can occur on a hanging rack as illustrated in FIG. 3 or ona work table as shown in FIG. 9. The stringer clip 26, shown in FIG. 10,is accurately positioned so that its laying surface 28 lines up exactlywith the corresponding vertical surface on the flange of the shear tie24 and both surfaces lie along the station plane around the fuselage asit will eventually exist from buildup of a plurality of panels 70.

While the skin 20 was on the fixture 34 a series of panel-to-panelcoordination holes 72 was drilled along the edge of the skin 22. Thesepanel-to-panel coordination holes 72 are now used to position the panelsrelative to each other on a fixture 74 shown in FIG. 11. Theconfiguration of the fixture 74 is non-critical because the panels arestill relatively flexible and the ultimate configuration of the fuselagewill be determined, not by the tooling, but by the parts themselves, aswill be described below.

The panels 70 are positioned on the fixture 74 and the coordinationholes 72 are aligned on adjacent panels and sealant is applied betweenthe facing surfaces of the panel edges. The panels are aligned so thatthe coordination holes 72 on adjacent panels line up exactly and the twopanels are fastened together at their adjacent edges by temporary clecofasteners through the coordination holes to insure that the panels areexactly aligned. The panels are then drilled and riveted to permanentlyfasten them together to form a super panel 71.

As shown in FIG. 12, an aircraft frame member 30 is mounted on a fixture76 in opposition to a precision machine tool 78 for drilling ofcoordination holes in the frame member. The frame is mounted on indexpins on index devices 82 similar to that shown in FIG. 7 and 8 and theposition of the index devices 82 is checked with the machine tool 78 asdiscussed below in connection with the checking of the machine tool 60.The position of the coordination holes is downloaded from the CAD/CAMmain frame original engineering part definition records in the samemanner that the machine tool 60 is controlled by original engineeringdata so that the coordination holes correspond to the original partdefinition rather than to an interpretation of that information asexpressed in hard tooling. The coordination holes drilled in the framemember 30 include holes 80 for alignment of the frame to thecoordination holes 69 in the stringer clips 26, holes 81 for alignmentof the floor grid as will be described later, and holes 83 for alignmentto the frame with stringer clips of the upper lobe of the fuselage, alsoto be described later.

After the coordination holes are drilled in the frame 30, the frame isremoved to the fixture 74 shown in FIG. 11 and fitted to the superpanel71. The coordination holes 80 are aligned with the coordination holes 69in the stringer clip 26 and are temporarily fastened through the alignedcoordination holes with cleco fasteners. This alignment determines theoutside contour of the superpanel skin so that the contour isindependent of any hard tooling and instead is determined by thelocation of the coordination holes 80 in the frame which in turn wasdirectly controlled by the original engineering part definition data asdownloaded from the data in the CAD/CAM main frame.

When the frame members 30 are all clecoed into position through thealigned coordination holes 80 and 69 in the frame and stringer clips 26,the tabs 84 on the frame member 30 are clamped, drilled and riveted tosecure the frame member 30 to the superpanel 71. The cleco fasteners arethen removed one by one and replaced with permanent rivets to secure thestringer clips 26 to the frame member 30 to produce a fuselage quarterpanel as shown in FIG. 13.

Turning now to FIG. 15, a floor grid 90 made of series of longitudinalfloor grid members 86 and cross floor grid members 88 is held in afixture 92 while coordination holes 94 are drilled in the ends of thecross members 88. The floor grid 90 itself may be manufactured using thesame determinant assembly concept of part manufacturing as disclosedherein for panel and fuselage manufacture. In this way, the longitudinaldimension between adjacent cross members 88 will correspond exactly tothe longitudinal dimension between adjacent station planes, so that whenfloor grid is inserted in the fuselage between frame members, the framemembers 30 will lie perfectly flush against the corresponding crossmembers 88 of the floor grid 90. When the coordination holes 94 havebeen drilled, the floor grid is hoisted by its support rack 96 intoposition in the assembled fuselage as shown in FIG. 16 and thecoordination holes 81 drilled by the machine tool 78 are aligned withthe coordination holes 94 in the ends of the cross members 88 of thefloor grid 90. Temporary cleco fasteners are inserted through thealigned coordination holes to hold the floor grid in its correctposition and the floor grid is then drilled and fastened into positionin permanent assembly as shown in FIG. 17. The fixture 98 on which thefuselage lobe 100 is supported has a compliant support 99 which allowsthe fuselage to flex so that the alignment of the coordination holes 81and 94 will determine the cross dimension of the fuselage lobe 100.Thus, the fixture 98 does not determine the fuselage contour, but ratherthe dimensions of the coordination holes drilled into the floor grid 90determines the cross dimension across the fuselage.

After all of the floor grid cross members 88 are fastened into place onthe frame members 30, an upper fuselage lobe 102, also made inaccordance with the determinate assembly technique described herein, ishoisted into place over the lower lobe 100 and coordination holes on theedges of the panels which make up the upper lobe are aligned with theholes in the panels which make up the lower lobe. Temporary clecofasteners are inserted through the aligned holes to hold the upper lobein place on the lower lobe. Coordination holes 83 in the protruding ends104 of the frame members are aligned with the corresponding stringerclips 26 in the upper lobe and are temporarily fastened with clecosthrough the aligned coordination holes. Once the frame members and lobeskin coordination holes are all aligned and temporarily fastened, theentire assembly is sealed, clamped, drilled and riveted to form thefinal fuselage section as shown in FIG. 18.

Turning now back to FIG. 5 and 6, the fuselage assembly cell 32 includesa monument 106 on which a pair of index devices 108 and 108' aremounted. The monument 106 can include a pair of fixed headers 110 and110' on which a test coupon can be mounted by way of index tabs whichfit into the index devices 108 and 108' in the same way that the skintabs 48 fit in the index device 40. The purpose of the test coupon is toprovide a test article that can be manufactured and accurately measuredin a convenient manner to provide assurance to the factory manager thatthe system is producing accurate parts.

A more important function of the monument is its role in an indexingsystem which is used each time a new operator assumes control over thefuselage assembly cell 32 from the previous operator at shift change orthe like. The new operator initiates a command which instructs themachine tool 60 to pick up a touch probe 112 and commence a probingroutine. The touch probe 112 can detect contact in the Z direction, thatis the direction of its length, and in both orthoginal directionsperpendicular to its length. This enables the touch probe 112 to measurethe distance from the machine tool to a contact point on thereconfigurable fixture 34 and also vertically and horizontally facingsurfaces on the reconfigurable fixture. The probe routine executed bythe robot arm first probes a pair of datum tings in the index devices108 and 108' attached to the monument 106 to determine any temperatureinduced changes between the two datum rings. A compensation factor isintroduced by the machine controller 68 to compensate for dimensionalchanges introduced by temperature deviations from the norm.

The touch probe 112 is guided by a custom program residing in themachine controller 68 to determine the location of the skin indexes 40and 40' mounted on the headers 36 of the reconfigurable fixture 34. Themeasured location is compared with the theoretical location and themachine controller 68 halts the operation if the actual location differsfrom the theoretical location by more than the specified tolerance.

A stationary probe 114 is mounted on one of the frame members of thereconfigurable fixture 34. The stationery touch probe 114 is used toperform a functional check of the end effectors by presenting the endeffectors to the stationery probe 114 to check for proper alignment andpart pickup. The robot arm 62 operating under a check routine, presentsthe end effector anvils to the stationery touch probe 114 and comparesthe measured results with the theoretical data as stored in the machinecontroller 68. If there is a deviation by more than the specifiedtolerance limits, the end effector may be replaced with a spare endeffector of the same type or the machine operation can be terminatedwhile an accuracy audit is conducted.

The spindle mounted probe 112 is used to determine the location of pointpositions along the curved surface of the headers 36. These pointlocations are transferred to the machine controller 68 for comparisonwith theoretical data to confirm that the sheet loaded on the fixture 34corresponds with the dataset loaded into the machine controller, andcorrective action is taken if deviations are noted. For example, anoffset amount may be established by which the data for the stringers,stringer clips, and shear ties can be offset when drilling and routingthe sheet.

Turning now to FIG. 20, the computing architecture for control of thefuselage assembly cell 32 is shown schematically to include computerfunctions which are performed by the CAD/CAM main frame 116 where theoriginal engineering digital product definition is recorded andavailable as the ultimate product definition authority. A numericalcontrol converts this data into a form that is usable by the postprocessor, which converts the digital parts definition data into motioncommands for the robot arm 62 when carrying the appropriate end effectorwith the appropriate cutter. The IMS Database is a large capacitystorage bank for storing all the parts programs that will be used by theassembly cell.

The other computing functions are performed at computer hardwarestations in the fuselage assembly cell 32 and for convenience areperformed by several separate computer hardware units, an IBM PS/2 118,the machine controller 68 which in the case of the preferred embodimentis an Allen Bradley 8600 IWS, and three Allen Bradley PLC's. One PLC 120controls the clamps in the end effector storage rack 64, and the othertwo PLC's 123 and 124 control the reconfigurable fixtures and man liftson the right and left banks of the assembly cell 32, respectively. Allthree PLC's communicate with the 8600 via a remote i/o line 126. The manlifts are personnel platforms 122 for raising workers up to theelevation of the skin index devices 40 when the skins are to be mountedon the reconfigurable fixture 34.

A system is thus disclosed which is usable for assembling parts on askin to produce an accurate fuselage panel, and a system is disclosedfor assembling panels so assembled into a full airplane fuselage. Thedeterminant assembly concept embodied in this disclosure utilizes thespatial relationships between key features of detail parts andsubassemblies, as defined in the digital design and represented bycoordination holes put into the parts and subassemblies by a numericallydriven tool controlled by original part design data, to control therelative location of detail parts in subassemblies, and the relativerelationship of subassemblies to each other, making the parts selflocating. This concept eliminates the need for traditional hard toolingused for decades in the air frame industry and for the first timeenables assembly of large mechanical structures wherein the contour ofthe structure and the relative dimensions within the structure aredetermined by the parts themselves rather than the tooling.

Obviously, numerous modifications and variations of the system disclosedherein will occur to those skilled in the art in view of thisdisclosure. Therefore, it is expressly to be understood that thesemodifications and variations, and the equivalents thereof, may bepracticed while remaining within the spirit and scope of the inventionas defined in the following claims, wherein:

We claim:
 1. A method of making a large airplane structure from aplurality of subassemblies, comprising:drilling coordination holes inselected components, which are made for accurate assembly of saidsubassemblies, and which will be located on said subassemblies inpositions to be used to accurately position said subassemblies relativeto each other so the spatial relationships between key features of saidsubassemblies, as defined by said coordination holes, make saidsubassemblies self-locating and intrinsically determinate of the finalcontour of said large airplane structure, independent of tooling;wherein said drilling of coordination holes in said selected componentsis done using an end effector carried by a precision computer controlledrobot that is directed to the drilling locations using a digital datasettaken directly from original digital part definition records; aligningsaid coordination holes in said selected components and assembling aplurality of subassemblies from said selected components; and assemblingsaid plurality of subassemblies to form said large airplane structure.2. The method as defined in claim 1, further comprising:checking theaccuracy of said robot by probing a monument of known dimensions andlocations to compare the known dimensions and location with thedimensions and location actually measured by said robot.
 3. The methodas defined in claim 1, further comprising:supporting said subassemblieson a compliant fixture that allows said large airplane structure to flexand assume a shape and dimensions as determined by the dimensions ofsaid subassemblies rather than the dimensions and shape of said fixture.